This invention relates generally to gas turbine engines and more particularly to the repair of turbine nozzle segments used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to a turbine section that extracts energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Gas turbine engines typically include stationary turbine nozzles that enhance engine performance by appropriately influencing gas flow and pressure within the turbine section. In multi-stage turbine sections, turbine nozzles are placed at the entrance of each turbine stage to channel combustion gases into the turbine rotor located downstream of the nozzle. Turbine nozzles are typically segmented around the circumference thereof with each nozzle segment having one or more airfoil vanes disposed between inner and outer bands that define the radial flowpath boundaries for the hot combustion gases flowing through the nozzle. These nozzle segments are mounted to the engine casing to form an annular array with the vanes extending radially between the rotor blades of adjacent turbine stages.
During operation, nozzle segments are exposed to a high temperature gas stream that can lead to oxidation and corrosion, thereby limiting the effective service life of these components. Accordingly, nozzle segments are typically fabricated from high temperature cobalt or nickel-based superalloys and are often coated with corrosion and/or heat resistant materials. Furthermore, nozzle segments (particularly those in the high pressure turbine section) are often cooled internally with cooling air extracted from the compressor to prolong service life. Even with such efforts, portions of the nozzle segments, particularly the vanes, can suffer parent metal cracking, material erosion due to oxidation and corrosion, and other damage such that the nozzle segments must be either repaired or replaced to maintain safe, efficient engine operation. Because nozzle segments are complex in design, are made of relatively expensive materials, and are expensive to manufacture, it is generally more desirable to repair them whenever possible.
One common repair process includes chemically stripping the environmental coating, applying a braze alloy to distressed areas to repair distress, and re-applying the environmental coating. However, such repair processes are limited by local distortion and under minimum wall thicknesses, which may be exceeded as a result of repeated repair and chemical stripping processes. That is, when the airfoil wall does not meet a minimum thickness, the nozzle segment cannot be repaired by the known repair process.
To avoid scrapping the entire nozzle segment in such situations, airfoil replacement techniques have been developed. Current airfoil replacement techniques for single vane segments comprise removing the distressed vane from the inner and outer bands and welding a new airfoil to the salvaged bands. When the distressed airfoil is removed from the bands, a stump of the original airfoil remains. The new airfoil is welded to these stumps. Airfoil welding works well with single vane segments because the required line-of-sight is available with the single vane configuration. However, conventional welding is not practical with multiple vane segments because of line-of-sight problems. Therefore, airfoil replacement for multiple vane segments typically comprises brazing new airfoils to the salvaged bands. This results in airfoil-to-band joints having generally inferior strength compared to that of an integrally cast interface.
Accordingly, it would be desirable to have an airfoil replacement technique for repairing multiple vane nozzle segments in which the airfoil-to-band joints are equivalent to or better than the same joints in an originally manufactured nozzle segment.
The above-mentioned need is met by the present invention, which provides a method of repairing a turbine nozzle segment having at least two vanes disposed between outer and inner bands. The method includes separating the outer and inner bands from the vanes, and forming one slot for each vane through the outer band and one slot for each vane through the inner band. A newly manufactured replacement vane is provided for each one of the original vanes. Each replacement vane has an outer boss formed on an outer end thereof and an inner boss formed on an inner end thereof. Then, each outer boss is inserted into a corresponding one of the outer band slots, and each inner boss is inserted into a corresponding one of the inner band slots. Once the parts are so assembled, each outer boss is welded to the outer band on the cold side, and each inner boss is welded to the inner band on the cold side. Each replacement vane is also brazed to the outer and inner bands on the respective hot sides thereof.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.